The present invention relates to materials designed to withstand high temperatures. More particularly, this invention relates to heat-resistant alloys for high-temperature applications, such as, for instance, gas turbine engine components of aircraft engines and power generation equipment.
There is a continuing demand in many industries, notably in the aircraft engine and power generation industries where efficiency directly relates to operating temperature, for alloys that exhibit sufficient levels of strength and oxidation resistance at increasingly higher temperatures. Gas turbine airfoils on such components as vanes and blades are usually made of materials known in the art as xe2x80x9csuperalloys.xe2x80x9d The term xe2x80x9csuperalloyxe2x80x9d is usually intended to embrace iron-, cobalt-, or nickel-based alloys, which include one or more additional elements to enhance high temperature performance, including such non-limiting examples as aluminum, tungsten, molybdenum, titanium, and iron. The term xe2x80x9cbasedxe2x80x9d as used in, for example, xe2x80x9cnickel-based superalloyxe2x80x9d is widely accepted in the art to mean that the element upon which the alloy is xe2x80x9cbasedxe2x80x9d is the single largest elemental component by weight in the alloy composition. Generally recognized to have service capabilities limited to a temperature of about 1100xc2x0 C. (about 2012xc2x0 F.), conventional superalloys used in gas turbine airfoils often operate at the upper limits of their practical service temperature range. In typical jet engines, for example, bulk average airfoil temperatures range between about 898xc2x0 C. (about 1650xc2x0 F.) to about 982 xc2x0 C. (about 1800xc2x0 F.), while airfoil leading and trailing edge and tip temperatures often reach about 1149xc2x0 C. (about 2100xc2x0 F.) or more. At such elevated temperatures, the oxidation process consumes conventional superalloy parts, forming a weak, brittle metal oxide that is prone to chip or spall away from the part. Maximum temperatures are expected in future applications to be over about 1315xc2x0 C. (about 2400xc2x0 F.), at which point many conventional superalloys begin to melt. Clearly, new materials must be developed if the efficiency enhancements available at higher operating temperatures are to be exploited.
The so-called xe2x80x9crefractory superalloys,xe2x80x9d as described in Koizumi et al., U.S. Pat. No. 6,071,470, represent a class of alloys designed to operate at higher temperatures than those of conventional superalloys. According to Koizumi et al., refractory superalloys consist essentially of a primary constituent selected from the group consisting of iridium (Ir), rhodium (Rh), and a mixture thereof, and one or more additive elements selected from the group consisting of niobium (Nb), tantalum (Ta), hafnium (Hf), zirconium (Zr), uranium (U), vanadium (V), titanium (Ti), and aluminum (Al). The refractory superalloys have a microstructure containing an FCC (face-centered cubic)-type crystalline structure phase and an L12xe2x80x94type crystalline structure phase, and the one or more additive elements are present in a total amount within the range of from 2 atom % to 22 atom %. As used herein, the term xe2x80x9crefractory superalloyxe2x80x9d is not limited to the definition of Koizumi et al., but it is used to refer to any alloy comprising a primary constituent (i.e., single largest constituent by weight) selected from the group consisting of rhodium and iridium, and further comprising an FCC (face-centered cubic)-type crystalline structure phase and an L12xe2x80x94type crystalline structure phase.
Although the refractory superalloys have shown potential to become replacements for conventional superalloys in present and future gas turbine engine designs, it has been shown that many alloys of this class do not meet all of the desired performance criteria for high-temperature applications. Therefore, the need persists for refractory superalloys with improved high-temperature properties.
The present invention provides several embodiments that address this need. One embodiment is an alloy comprising from about three atomic percent to about nine atomic percent of at least one precipitation-strengthening metal selected from the group consisting of zirconium, niobium, tantalum, titanium, hafnium, and mixtures thereof; from about one atomic percent to about five atomic percent ruthenium (Ru); and the balance rhodium; the alloy further comprising a face-centered-cubic phase and an L12xe2x80x94structured phase.
A second embodiment is a gas turbine engine component comprising an alloy, the alloy of the component comprising from about three atomic percent to about nine atomic percent of at least one precipitation-strengthening metal selected from the group consisting of zirconium, niobium, tantalum, titanium, hafnium, and mixtures thereof; from about one atomic percent to about five atomic percent ruthenium; and the balance rhodium; the alloy of the gas turbine engine component further comprising a face-centered-cubic phase and an L12xe2x80x94structured phase.